National Advisory Committee for Aeronautics N0.91 R esearc h Abstracts OCTOBER 21, 1955 CURRENT NACA REPORTS NACA Rept. 1191 ON THE DEVELOPMENT OF TURBULENT WAKES FROM VORTEX STREETS. Anatol Roshko, Cali- fornia Institute of Technology. 1954. ii, 25p. diagrs., photos., 3 tabs. (NACA Rept. 1191. Formerly TN 2913) Wake development behind circular cylinders at Reynolds numbers from 40 to 10,000 was investigated by hot-wire techniques in a low-speed wind tunnel. The Reynolds, number range of periodic vortex shedding is divided into two distinct subranges. In the stable range, R = 40 to 150, regular vortex streets are formed and no turbulent motion develops, the vortices decaying by viscous diffusion. The range.R = 150 to 300 is a transition region to the irregular range in which turbulent velocity fluctua- tions accompany the periodic formation of vortices. The diffusion is turbulent and the wake becomes fully turbulent in 40 to 50 diameters. The turbulence is initiated by laminar-turbulent transition in the free layers which spring from the separation points on the cylinder. An annular vortex street was ob- served in the wake of a ring. NACA RM E55E18 PERFORMANCE CHARACTERISTICS OF HEMI- SPHERICAL TARGET-TYPE THRUST REVERSERS. Fred W. Steffen, Jack G. McArdle, and James W. Coats. September 1955. 39p. diagrs., photos., tab. (NACA RM E55E18) Reverse-thrust performance of hemispherical target -type thrust reversers was obtained. The value of reverse-thrust ratio was found to be pri- marily a function of hemisphere diameter. Slight improvements in performance were obtained with some size reversers by operating with the exhaust nozzle opened slightly. When high values of reverse-thrust ratio were obtained, the reversed flow attached to the boattail and the pressures on the boattail fell belovv atmospheric. This pressure re- duction amounted to about 20 percent of the resultant reverse force. Various boattail shapes had little ef- fect on reverse-thrust ratio for most hemisphere sizes. NACA RM E55G27a IDEAL TEMPERATURE RISE DUE TO CONSTANT- PRESSURE COMBUSTION OF A JP-4 FUEL. S(idney) C. Huntley. September 1955. 53p. diagrs., tabs. (NACA RM E55G27a) Charts are presented from which ideal temperature rise or the ideal quantity of fuel required to obtain a specified combustion temperature may be obtained. The charts are applicable only to a fuel having a hydrogen-carbon mass ratio of 0.168 (CH2) and in- clude a range of fuel-air ratios from to 1.2 frac- tions of stoichiometric fuel-air ratio >vith dissocia- tion taken into account, inlet-air temperatures from 400° to 1600° R, and combustion pressures from' 1/16 to 64 atmospheres NACA RM L55G21 NOTE ON HOVERING TURNS WITH TANDEM HELICOPTERS. John P. Reeder and Robert J. Tapscott. September 1955. 5p. photo. (NACA RM L55G21) The source of an appreciable pitching-moment dif- ference between left and right hovering turns for a tandem helicopter is described. The difference in pitching moment results from the difference in ro- tational speed of the counterrotating rotors with respect to the air while the helicopter is turning. NACA TM 1377 THE THEORIES OF TURBULENCE. (Les Theories de la Turbulence). L. Agostini and J. Bass. October 1955. 163p. diagrs. (NACA TM 1377. Trans, from Ministere de l'Air, Publications Scientifiques et Techniques 237, 1950) The report includes a discussion of the kinematics of statistical mediums, particularly those which are isotropic. A mathematical study is made of the ap- plications of Navier's equations to turbulent motion. Physical theories involving similarity are dealt with. Review is made of much of the work in turbulence. The theoretical discussions are illustrated by some correlation and spectrum curves based on measure- ments taken in the wind tunnel at the laboratory of the mechanics of the atmosphere at Marseille. U.S. DEPOSITORY } •AVAILABLE ON LOAN ONLY, ADDRESS REQUESTS FOR DOCUMENTS TO NACA, 1512 H ST., NW„ WASHINGTON 25, D C ., CITING CODE NUMBER ABOVE EACH TITLE; THE REPORT TITLE AND AUTHOR. 2.% /3CX? NACA TM 1389 OPTIMUM FLIGHT PATHS OF TURBOJET AIR- CRAFT. (Traiettorie Ottime Di Volo Degli Aero- plani Azionati Da Turboreattori). Angelo Miele. September 1955. 47p. diagrs. , tabs. (NACA TM 1389. Trans, from L'Aerotecnica, v. 32, no. 4, 1952, p. 206-219) The climb of turbojet aircraft is analyzed and dis- cussed including the effects of tangential accelera- tions. Three flight performances are examined: minimum time of climb, climb with minimum fuel consumption, and steepest climb. Diagrams for quick calculation of the optimum Mach numbers and the effect of acceleration on the rate of climb in tropospheric and stratospheric flight are given. NACA TM 1396 FROM LINEAR MECHANICS TO NONLINEAR MECHANICS. (De la mecanique lineaire a la mecanique non lineaire). Julien Loeb. October 1955. 18p. diagrs.^ ^NACA TM 1396. Trans, from Annales des Telecommunications, v. 5, no. 2, Feb., 1950, p. 65-71) Consideration is first given to the technique used in telecommunication where a nonlinear system (the modulator) results in a linear transposition of a signal. It is then shown that a similar method per- mits linearization of electromechanical devices or nonlinear mechanical devices. A sweep function plays the same role as the carrier wave in radio- electricity. The linearizations of certain non- linear functionals are presented. NACA TN 3463 INVESTIGATION OF THE VIBRATIONS OF A HOL- LOW THIN-WALLED RECTANGULAR BEAM. Eldon E. Kordes and Edwin T. Kruszewski. October 1955. 24p. diagrs., photos., 2 tabs. (NACA TN 3463) Experimental modes and frequencies of an unstif- fened hollow beam of rectangular cross section are presented, and comparisons are made between ex- perimental and theoretical frequencies. Theories based on rigid cross sections were found to be suf- ficiently accurate to predict the frequencies of only the lovvest three bending modes. For the higher bending modes and all the torsional modes, it was necessary to include the effects of cross-sectional distortions in the calculations. NACA TN 3464 INFLUENCE OF SHEAR DEFORMATION ON THE CROSS SECTION ON TORSIONAL FREQUENCIES OF BOX BEAMS. Edwin T. Kruszewski and William W. Davenport. October 1955. 23p. diagrs. (NACA TN 3464) An exact analysis has been carried out on the tor- sional vibrations of a four-flange box beam with NACA RESEARCH ABSTRACTS NO. 91 cross sections which can change shape because the stiffness of the bulkheads is finite. The effect of shear deformation of the cross section on the tor- sional frequencies is illustrated by numerical cal- culations. An approximate method for quickly estimating the effects of bulkhead shear stiffness on the torsional frequencies of box beams has been devised. NACA TN 3475 AN ANALYSIS OF ACCELERATION, AIRSPEED, AND GUST-VELOCITY DATA FROM ONE TYPE OF FOUR-ENGINE TRANSPORT AIRPLANE OPERATED OVER TWO DOMESTIC ROUTES. Martin R. Copp and Thomas L. Coleman. October 1955. 29p. diagrs., tabs. (NACA TN 3475) Time-history data obtained by the NACA VGH re- corder from one type of four-engine commercial transport airplane during operations on two domestic routes indicated that the number of gust accelera- tions experienced per mile of flight by the two opera- tions differed by a factor of roughly 3. The number of gusts per mile of flight differed by a factor of roughly 4. A general decrease in the frequency of occurrence of gust velocities with increasing altitude was noted for both operations. For acceleration values above 0.8g, maneuver accelerations formed a substantial part of the total-flight load histories. A comparison of the average overall airspeeds and the corresponding average airspeeds in rough air (with accelerations equal to or greater than 0.3g) indicated very little slowdown by either operation upon en- countering turbulence. NACA TN 3481 WIND-TUNNEL INVESTIGATION AT LOW SPEED OF EFFECT OF SIZE AND POSITION OF CLOSED AIR DUCTS ON STATIC LONGITUDINAL AND STATIC LATERAL STABILITY CHARACTERISTICS OF UNSWEPT -MID WING MODELS HAVING WINGS OF ASPECT RATIO 2, 4, AND 6. Byron M. Jaquet and James L. Williams. September 1955. 45p. diagrs., photos., tabs. (NACA TN 3481) Results are presented of tests made at a Mach num- ber of 0.13 in the Langley stability tunnel to deter- mine the effects of closed wing-root air ducts on the static longitudinal and static lateral stability charac- teristics of models having unswept wing and tail sur- faces with wings of aspect ratio 2, 4, and 6. In addition, for model configurations employing the wing of aspect ratio 2 the effects of top and bottom fuselage ducts on the static longitudinal and static lateral characteristics were determined. The effect of the wing-root ducts on the aerodynamic hystere- sis in sideslip of the model employing the wing of aspect ratio 2 was also determined. NACA RESEARCH ABSTRACTS NO. 91 NACA TN 3485 AN APPROXIMATE SOLUTION FOR AXIALLY SYMMETRIC FLOW OVER A CONE WITH AN AT- TACHED SHOCK WAVE. Richard A. Hord. October 1955. 32p. diagrs. (NACA TN 3485) It is shown that the streamlines in an angular neighborhood of the surface of an unyawed circular cone with an attached shock wave are, to a first ap- proximation, portions of hyperbolas. This fact is used as a basis for the development of an approxi- mate solution in which shock-wave orientation and flow field behind the shock wave are given explicitly in terms of free-stream Mach number, vertex angle of the body cone, and the ratio of specific heats of the gas. The approximate solution is compared with other approximate solutions for the cone. NACA TN 3497 SUMMARY OF RESULTS OF A WIND-TUNNEL INVESTIGATION OF NINE RELATED HORIZONTAL TAILS. Jules B. Dods, Jr. and Bruce E. Tinling. July 1955. 105p. diagrs., 2 tabs. (NACA TN 3497. Formerly RM A51G31a) A compilation of data is presented for models of nine related horizontal tails. The majority of the results were obtained at a Mach number of approxi- mately 0.20. Three of the models were tested throughout the subsonic Mach number range to a maximum of 0.94. The Reynolds number range was from 2 to 4 million. The models had aspect ratios from 2 to 6, angles of sweepback from 5. 7° to 45°, and had 30-percent-chord, sealed, plain flaps. The lift coefficient, hinge-moment coefficient, and pres- sure coefficients across the elevator nose seal are presented. The effects of sweepback, aspect ratio, and Mach number on the lift and hinge-moment parameters are summarized. Comparisons of the experimental results with theoretical calculations are presented. NACA TN 3519 VISUALIZATION STUDY OF SECONDARY FLOWS IN TURBINE ROTOR TIP REGIONS. Hubert W. Allen and Milton G. Kofskey. September 1955. 33p. diagrs., photos., tab. (NACA TN 3519) A low-speed visualization study of turbine rotor tip secondary flows was made. Results include qualita- tive information on tip-clearance flow, cross- passage flow, and scraping flow, and on a range of rotor speeds for which a transition condition appear- ed with minimum flow disturbance. Rotor speed re- quired for this transition flow condition depended upon blade configuration, angle of incidence, tip clearance, and air-flow rate. Results should aid in extending the study to higher airspeeds. NACA TN 3532 LOW-SPEED STATIC LATERAL AND ROLLING STABILITY CHARACTERISTICS OF A SERIES OF CONFIGURATIONS COMPOSED OF INTERSECTING TRIANGULAR PLAN-FORM SURFACES. David F. Thomas, Jr. October 1955. 29p. diagrs., photos. (NACA TN 3532) The static lateral and rolling stability derivatives of a series of cruciform, inverted T-, V-, and Y- configurations composed of low-aspect-ratio trian- gular surfaces have been obtained at low speed in the 6-foot-diameter rolling-flow test section of the Langley stability tunnel. These derivatives are presented as functions of the geometry of the models, and for two configurations (a planar wing and an in- verted T), as functions of angle of attack. Where possible, comparisons have been made to indicate the extent of agreement between experiment and existing theory. In general, the sideslip deriva- tives showed better agreement between theory and experiment than the rolling derivatives. NACA TN 3533 THE PROPER COMBINATION OF LIFT LOADING FOR LEAST DRAG ON A SUPERSONIC WING. Frederick C. Grant. October 1955. 21p. diagrs. (NACA TN 3533) Lagrange's method of undetermined multipliers is applied to the problem of properly combining lift loadings for the least drag at a given lift on super- sonic wings. The interference drag between the optimum loading and any loading at the same lift coefficient is found to be constant on a given plan form. This is an integral form of a criterion established by Robert T. Jones for optimum load- ings. The best combination of four loadings on a delta wing with subsonic leading edges is calculated for several Mach numbers as a numerical example. NACA TN 3535 FLIGHT INVESTIGATION OF THE SURFACE- PRESSURE DISTRIBUTION AND THE FLOW FIELD AROUND A CONICAL AND TWO SPHERICAL NON- ROTATING FULL-SCALE PROPELLER SPINNERS. Jerome B. Hammack, Milton L. Windier, and Elwood F. Scheithauer. September 1955. 36p. diagrs. , photos. (NACA TN 3535) The surface-pressure distribution and the flow field around a conical and two spherical nonrotating full- scale propeller spinners were determined in flight at Mach numbers of 0. 70 to 0.96. The local-surface Mach numbers between the cone-sphere tangency and the maximum thickness stations were approximately 0.4 higher than those at a free-stream Mach number of 0.95. The departure from free-stream conditions of the larger spherical spinner extends beyond the 1. 3-spinner-radius station; whereas, with the smaller spherical spinner, free-stream conditions were reached at the 1. 3-spinner-radius station. NACA TN 3536 A LIMITED FLIGHT INVESTIGATION OF THE EF- FECT OF THREE VORTEX-GENERATOR CONFIGU- RATIONS ON THE EFFECTIVENESS OF A PLAIN FLAP ON AN UNSWEPT WING. Garland J. Morris and Lindsay J(ohn) Lina. September 1955. 20p. diagrs. , photos. , tabs. (NACA TN 3536) An exploratory flight investigation was made to deter- mine the effect of three vortex-generator configura- tions on the effectiveness of the plain flap of a fighter airplane. Tests were made with flaps deflected 19° and 45° at several indicated airspeeds in the range from stall to 140 miles per hour. No improvement in flap effectiveness was obtained with the flaps de- flected full down. With flaps deflected 19°, the vortex-generator configuration which produced the largest increase in lift coefficient had about the same effect as increasing the flap deflection to about 27°. Another configuration produced some increase in lift coefficient with no apparent increase in drag. NACA TN 3538 SUMMARY OF DERIVED GUST VELOCITIES OB- TAINED FROM MEASUREMENTS WITHIN THUN- DERSTORMS. H(arold) B. Tolefson. October 1955. 19p. diagrs., tabs. (NACA TN 3538) Available measurements of the derived gust veloci- ties within thunderstorms are summarized for alti- tudes from 5,000 to 34,000 feet. The results indi- cate that the intensity of the derived gust velocity is essentially constant up to altitudes of 20,000 feet and that an approximate 10-percent reduction in the gust intensity occurs for altitudes from 20,000 to 30,000 feet. NACA TN 3561 INTENSITY, SCALE, AND SPECTRA OF TURBU- LENCE IN MIXING REGION OF FREE SUBSONIC JET. James C. Laurence. September 1955. 58p. diagrs., photo., tab. (NACA TN 3561) Hot-wire anemometer measurements of the turbu- lence parameters were made in a 3.5-inch-diameter free jet at exit Mach numbers between 0.2 and 0.7 and Reynolds numbers (based on jet rad. ) between 37,500 and 350,000. The results of these measure- ments show that (1) the intensity of turbulence is a max. at a distance of approximately 1 jet rad. from the jet center line and decreases with increasing Mach and/or Reynolds number, and (2) the lateral and longitudinal scales of turbulence are nearly in- dependent of Mach and/or Reynolds number and vary proportionally with distance from the jet nozzle. The lateral scale is much smaller than the longi- tudinal and does not vary with distance from the center line, while the longitudinal scale is a max. at a distance from the center line of about 0.7 to 0.8 of the jet rad. NACA RESEARCH ABSTRACTS NO. 91 NACA TN 3568 AVERAGING OF PERIODIC PRESSURE PULSA- TIONS BY A TOTAL-PRESSURE PROBE. R. C. Johnson. October 1955. 30p. diagrs., photo., tabs. (NACA TN 3568) Information is presented on the average pressure indicated by a total -pressure probe subjected to a stagnation pressure that alternates periodically be- tween two constant values. Calculated and experi- mental data are in good agreement, and errors are reduced when the probe design is such as to ensure laminar-flow pulsations in the probe at all times. The averaging error is minimized when the inside diameter of the probe entrance tube is made as small as possible, and its length as great as possi- ble, consistent with an acceptable time lag. NACA TN 3570 AN EXPERIMENTAL COMPARISON OF THE LAGRANGIAN AND EULERIAN CORRELATION COEFFICIENTS IN HOMOGENEOUS ISOTROPIC TURBULENCE. William R. Mickelsen. October 1955. 42p. diagrs. (NACA TN 3570) The Lagrangian and Eulerian correlation coeffi- cients were compared in a field of homogeneous, isotropic turbulence. The Lagrangian correlation coefficient was characterized by diffusion measure- ments, and the Eulerian coefficient was measured by hot-wire anemometry. The Lagrangian and Eulerian correlation coefficients had similar shapes connected by a linear relation between their coordi- nates. The proportionality factor in the linear rela- tion was roughly constant over a range of turbulence intensities from 1.8 to 14 feet per second. The lin- ear relation permits solution of mixing problems from the Eulerian turbulence parameters. BRITISH REPORTS N-38869* Royal Aircraft Establishment (Gt. Brit.) A SIMPLIFIED MODEL OF THE INCOMPRESSIBLE FLOW PAST TWO-DIMENSIONAL AEROFOILS WITH A LONG BUBBLE TYPE OF FLOW SEPARA- TION. J. F. Norbury and L. F. Crabtree. June 19,55. 17p. diagrs., tabs. (RAE Tech. Note Aero 2352) This problem may be divided into two parts. The first part concerns the external inviscid flow, and the other is related to the details of the flow inside the viscous region formed by the bubble, the bound- ary layer, and the wake. Although much informa- tion can be obtained by considering only the external stream (as in the hodograph method developed by Maskell), a complete and unique solution for the pressure distribution can only be found by attacking the second part of the problem as well. This has been attempted here and it is shown how the pres- sure recovery ratio at the end of a long bubble may be found by the analysis of a simplified model of the flow; the complete pressure distribution may then be uniquely determined. NACA RESEARCH ABSTRACTS NO. 91 MISCELLANEOUS N-39013* AIRCRAFT STRUCTURES RESEARCH AT ELEVAT- ED TEMPERATURES. John E. Duberg. (Present- ed to Structures and Materials Panel of Advisory Group for Aeronautical Research and Development (NATO) London, England, September 5-9, 1955). 36p. diagrs. , photos. A review is made of the test techniques that have been developed and used by the NACA for experimen- tal research in aircraft structures at elevated temp- eratures. Some experimental results are presented. Remarks are included on the problem of model scal- ing for testing of structures at high temperatures. UNPUBLISHED PAPERS N-37996* THE BUCKLING OF PLATES AND BARS IN THE PLASTIC RANGE. I. THEORY. (Over het knitvraagstuk in het plastische gebied bij staven en platen. I. Theorie). J. F. Besseling. July 1955. 97p. diagrs., tabs. (Trans, from Nationaal Luchtvaartlaboratorium, Amsterdam, S.407, Oct. 14, 1952) The topics treated include the plastic buckling of bars and plates; application of plasticity theories to 'the determination of stiffness quantities needed for computing the buckling; and the effect of plasticity on the buckling load for a number of fundamental buckling cases when the buckling stress condition is reached through a uniform increase of loading. This report is limited to the theoretical side of the buck- ling problem and gives the basis for comparing theoretical and experimental results. DECLASSIFIED NACA REPORTS NACA RM A8I17 INVESTIGATION OF A THIN WING OF ASPECT RATIO 4 IN THE AMES 12-FOOT PRESSURE WIND TUNNEL. Ill - THE EFFECTIVENESS OF A CONSTANT-CHORD AILERON. Ben H. Johnson, Jr. and Fred A. Demele. November 19, 1948. 26p. diagrs., photo. (NACA RM A8I17) (Declassified from Confidential, 9/15/55) Presented in the report are results of tests at Mach numbers from 0.27 to 0.94 of a thin, unswept wing having a modified diamond airfoil section of thickness ratio 0.042 and equipped with a constant-chord ailer- on. The tests were conducted at a constant Reynolds number of 2, 730, 000. The effects of compressibility on the aileron effectiveness were negligible at Mach numbers up to 0.85, but at higher Mach numbers er- ratic effects of compressibility were evident, espe- cially at lift coefficients greater than 0.5. NACA RM A9I01 INVESTIGATION OF A THIN WING OF ASPECT RATIO 4 IN THE AMES 12-FOOT PRESSURE WIND TUNNEL. V - STATIC LONGITUDINAL STABILITY AND CONTROL THROUGHOUT THE SUBSONIC SPEED RANGE OF A SEMISPAN MODEL OF A SUPERSONIC AIRPLANE. Ben H. Johnson, Jr. . and Francis W. Rollins. December 8, 1949. 130p. diagrs., photos. (NACA RM A9I01) (Declassified from Confidential, 9/15/55) Presented in this report are results of tests at Mach numbers from 0.20 to 0.94 of a model of a hypo- thetical supersonic airplane equipped with a thin, sharp-edged wing and tail without sweep. The static longitudinal-stability characteristics of the model have been measured for two different vertical loca- tions of the horizontal tail. The longitudinal control afforded by an all-movable stabilizer and by an ele- vator has been investigated. The downwash at the tail has been computed and the dynamic-pressure ratio at the tail has been evaluated from pressure measurements in the wake of the wing. NACA RM A51I07 THE STATIC LONGITUDINAL CHARACTERISTICS AT MACH NUMBERS UP TO 0.95 OF A TRIANGULAR-WING CANARD MODEL HAVING A TRIANGULAR CONTROL. Jack D. Stephenson and Ralph Selan. December 1951. 72p. diagrs., photo. (NACA RM A51I07) (Declassified from Confidential, 9/15/55) Presents and analyzes results of tests to assess the longitudinal characteristics of a canard-type model having a triangular wing of aspect ratio 2 and NACA 0008-63 sections. The horizontal canard surface had a plan form identical to the wing and an NACA 0005-63 section. The Mach number range of the investigation varied from 0.25 to 0.95 at Reynolds numbers of 8 million and 3 million, respectively. The model was tested with the horizontal control surface at various fixed angles of incidence and with the surface unrestrained so that it could pivot about an axis at 30 percent of its mean aerodynamic chord. NACA RM A52F18 THE LONGITUDINAL CHARACTERISTICS AT MACH NUMBERS UP TO 0.92 OF A CAMBERED AND TWISTED WING HAVING 40° OF SWEEPBACK AND AN ASPECT RATIO OF 10. George G. Edwards, Bruce E. Tinling, and Arthur C. Ackerman. September 1952. 71p. diagrs., photos., tab. (NACA RM A52F18) (Declassified from Confidential, 9/15/55) A sweptback wing, in combination with a fuselage, of a type considered suitable for long-range, high- speed airplanes, has been tested in the Ames 12-foot pressure wind tunnel. The wing had 40° of sweep- back, an aspect ratio of 10, a taper ratio of 0.4, and 5° of washout at the tip. The sections normal to the sweep reference line had NACA four-digit profiles, design lift coefficients of 0.40, and varied in thick- ness ratio from 14 percent at the root to 11 percent at the tip. The lift, drag, and pitching moment of a semispan model were measured at Reynolds num- bers from 2,000,000 to 8,000.000 at low Mach num- bers, and at Mach numbers from 0.25 to 0.92 at a Reynolds number of 2,000,000. The changes in boundary-layer flow on the upper surface were studied with tufts. NACA RM A52I19 THE LONGITUDINAL CHARACTERISTICS AT MACH NUMBERS UP TO 0.9 OF A WING-FUSELAGE-TAIL COMBINATION HAVING A WING WITH 40° OF SWEEPBACK AND AN ASPECT RATIO OF 10. Bruce E. Tinling. December 1952. 41p. diagrs., photo., tab. (NACA RM A52I19) (Declassified from Confidential, 9/15/55) Wind-tunnel tests were made of a wing-fuselage hori- zontal tail combination suitable for long-range high- speed airplanes. The cambered and twisted wing had an aspect ratio of 10, taper ratio of 0.4, and 40° sweepback. The all-movable horizontal tail had an aspect ratio of 4.5, taper ratio of 0.4, and 40° sweepback. Included are data on longitudinal sta- bility and control, wing fence and tail effect, and isolated horizontal tail data. Data »vere obtained at a Reynolds number of 8,000,000 at low Mach num- bers, and at a Reynolds number of 2,000,000 at Mach numbers up to 0.9. NACA RM A52K20 PRESSURE DISTRIBUTION AT MACH NUMBERS UP TO 0.90 ON A CAMBERED AND TWISTED WING HAVING 40° OF SWEEPBACK AND AN ASPECT RATIO OF 10, INCLUDING THE EFFECTS OF FENCES. Frederick W. Boltz and Harry H. Shibata. March 1953. 133p. diagrs., photos., tabs. (NACA RM A52K20) (Declassified from Confidential, 9/15/55) Pressure-distribution measurements have been made on a semispan model of a cambered and twisted wing, NACA RESEARCH ABSTRACTS NO. 9) alone and in combination with a fuselage. The wing had 40° of sweepback, an aspect ratio of 10, a taper ratio of 0.4, and 5° of washout at the tip. The wing thickness distribution in sections normal to the ref- erence sweep line was the NACA 4-digit series with the maximum thickness varying from 14-percent chord at the root to 11-percent chord at the tip. The chordwise distributions of pressure coefficient at nine semispan stations are presented for Mach num- bers of 0.165 and 0.25 at a Reynolds number of 8,000,000, and for Mach numbers from 0.25 to 0.90 at a Reynolds number of 2,000,000. Tabulated pressure data are presented for the wing without fences and with four fences. NACA RM A53D06 THE EFFECTS OF NACELLES AND OF EXTENDED SPLIT FLAPS ON THE LONGITUDINAL CHARAC- TERISTICS OF A WING-FUSELAGE-TAIL COMBI- NATION HAVING A WING WITH 40° OF SWEEP- BACK AND AN ASPECT RATIO OF 10. Bruce E. Tinling and Armando E. Lopez. June 1953. 47p. diagrs., tab. (NACA RM A53D06) (Declassified from Confidential, 9/15/55) Wind-tunnel tests were made to evaluate the effects of nacelles and of extended split flaps on the longitu- dinal characteristics of a wing-fuselage-tail combi- nation suitable for long-range high-speed airplanes. The cambered and twisted wing had an aspect ratio of 10, a taper ratio of 0.4, and 40° of sweepback. The nacelles were at 25 and 50 percent of the semi- span. Data were obtained to evaluate the effects of nacelles at Mach numbers up to 0.90 at a Reynolds number of 2,000,000. The effects of the flaps were evaluated from data obtained at a Reynolds number of 4,000,000 and a Mach number of 0.082. A limited number of data were also obtained to evalu- ate the effects of increasing the tail height. NACA RM A53I23 DOWNWASH BEHIND A TRIANGULAR WING OF ASPECT RATIO 3 - TRANSONIC BUMP METHOD. John A. Axelson. December 1953. 37p. diagrs., photo., tab. (NACA RM A53I23) (Declassified from Confidential, 9/15/55) Downwash measured by means of an all-movable horizontal tail in several different locations is pre- sented for a triangular wing having an aspect ratio of 3 and the NACA 63A006 section. The lift, drag, and pitching-moment characteristics of the wing and the downwash are presented for angles of attack up to 28° over a Mach number range from 0.6 to 1.1, corresponding to a Reynolds number range from 1.8 million to 2.4 million. The effects of tail location on the tail contribution to the static longitudinal sta- bility of the wing-tail combination are discussed. NACA RESEARCH ABSTRACTS NO. 91 NACA RM A53I28 THE RESULTS OF WIND-TUNNEL TESTS AT LOW SPEEDS OF A FOUR-ENGINE PROPELLER-DRIVEN AIRPLANE CONFIGURATION HAVING A WING WITH 40° OF SWEEPBACK AND AN ASPECT RATIO OF 10. George G. Edwards, Donald A. Buell, and Jerald K. Dickson. December 1953. 121p. diagrs., photo., tabs. (NACA RM A53I28) (Declassified from Confidential, 9/15/55) The effects of operating propellers on the low-speed longitudinal characteristics of a four-engine tractor airplane configuration having a sweptback wing have been investigated in the Ames 12-foot pressure wind tunnel at thrust coefficients up to 0.9 per propeller and at Reynolds numbers from 4,000,000 to 8,000,000. Variations in the model included different heights and incidences of the horizontal tail as well as tail removed, two arrangements of extended split flaps, several propeller-blade angles, and inde- pendent as well as simultaneous operation of the in- board and outboard propellers. Coefficients of lift, longitudinal force, pitching moment, propeller thrust, and propeller power are presented in tabular form for various values of advance ratio at constant angles of attack. Selected portions of the data are presented in plotted form for various constant thrust coefficients. NACA RM A53I29 THE TRANSONIC CHARACTERISTICS OF 36 SYM- METRICAL WINGS OF VARYING TAPER, ASPECT RATIO, AND THICKNESS AS DETERMINED BY THE TRANSONIC -BUMP TECHNIQUE. Warren H. Nelson, Edwin C. Allen, and Walter J. Krumm. December 1953. 131p. diagrs., photo. (NACA RM A53I29) (Declassified from Confidential, 9/15/55) An investigation was made in the Ames 16-foot high- speed wind tunnel, utilizing the transonic-bump tech- nique to determine the effects of plan-form taper on a series of wings having aspect ratios of 4, 3, and 2, and NACA 63A00X sections with thickness-to-chord ratios of 8, 6, 4, and 2 percent. The Mach number range was 0.6 to 1.1 with a corresponding Reynolds number range of about 1.4 million to 2.0 million. The results indicate that increasing the taper ratio caused only small increases in lift-curve slope except for the wings of highest aspect ratio and thinnest sections. Increasing taper ratio generally increased the overall center-of-pressure travel in going from subsonic to supersonic speeds. NACA RM E8I01 EFFECT OF THREE FLAME -HOLDER CONFIGU- RATIONS ON SUBSONIC FLIGHT PERFORMANCE OF RECTANGULAR RAM JET OVER RANGE OF ALTITUDES. Dugald O. Black and Wesley E. Messing. November 24, 1948. 28p. diagrs., photo., tab. (NACA RM E8I01) (Declassified from Confidential, 9/15/55) A flight investigation is reported that was conducted on a rectangular ram jet incorporating various flame-holder configurations over a range of fuel -air ratios from 0.017 to 0.120, combustion-chamber- inlet velocities from 50 to 125 feet per second, and pressure altitudes from 1500 to 28,000 feet. Highest combustion efficiencies, which varied from a maxi- mum of 82 percent at 1500 feet to 39 percent at 26,000 feet, were obtained at all altitudes with the three-V gutter-type flame holder. However, at any given altitude and fuel-air ratio slightly higher net- thrust coefficients occurred with the two-V flame holder as a result of its lower value of pressure drop. NACA RM E8I28 EFFECT OF VARIATION IN FUEL PRESSURE ON COMBUSTION PERFORMANCE OF RECTANGULAR RAM JET. Wesley E. Messing and Dugald O. Black. November 24, 1948. 26p. diagrs., photo., tab. (NACA RM E8I28) (Declassified from Confidential, 9/15/55) Reports effect of variation in fuel pressure on start- ing characteristics, minimum blow-out limits, com- bustion efficiencies, gas total-temperature ratio, and net -thrust coefficient of rectangular ram jet operated over range of pressure altitudes from 1500 to 26, 300 feet, indicated airspeeds from 100 to 200 miles per hour, and fuel-air ratios from 0.017 to 0.120. In general, increasing the degree of fuel atomization and distribution by utilization of small orifice fuel nozzles that operated at high fuel pressures resulted in higher values of combustion efficiency, gas total- temperature ratio, and net-thrust coefficient at a given fuel -air ratio. NACA RM E50D28 AERODYNAMIC CHARACTERISTICS OF NACA RM- 10 MISSILE IN 8- BY 6-FOOT SUPERSONIC WIND TUNNEL AT MACH NUMBERS FROM 1.49 TO 1.98. H - PRESENTATION AND ANALYSIS OF FORCE MEASUREMENTS. Fred T. Esenwein, Leonard J. Obery, and Carl F. Schueller. July 21, 1950. 34p. diagrs., photo. (NACA RM E50D28) (Declassified from Confidential, 9/15/55) Experimental investigation of aerodynamic forces acting on body of revolution (NACA RM-10 missile) with and without stabilizing fins was conducted at Mach numbers from 1.49 to 1.98 at angles of attack from 0° to 9° and at Reynolds number of approxi- mately 30, 000, 000. Comparison of experimental lift, drag, and pitching-moment coefficients and center-of-pressure location for body alone is made with linearized potential theory and a semiempirical method. Results indicate that aerodynamic charac- teristics were predicted more accurately by semi- empirical method than by potential theory. Break- down of measured drag coefficients into components of friction, pressure, and base-pressure drag is presented for body alone at zero angle of attack. NACA RM E 5011 9 AERODYNAMIC CHARACTERISTICS OF NACA RM-10 MISSILE IN 8- BY 6-FOOT SUPERSONIC WIND TUNNEL AT MACH NUMBERS FROM 1.49 TO 1.98. m - ANALYSIS OF FORCE DISTRIBUTION AT ANGLE OF ATTACK (STABILIZING FINS RE- MOVED). Roger W. Luidens and Paul C. Simon. December 12, 1950. 26p. diagrs. (NACA RM E50I19) (Declassified from Confidential 9/15/55) Analysis of force distribution on slender pointed body of revolution at angle of attack was made utiliz- ing pressure-distribution data and balance measure- ments obtained in NACA Lewis 8- by 6-foot super- sonic tunnel. Comparison of experimental station normal force with those predicted by linearized po- tential theory (on which radius of body assumed to approach zero) shows that inaccurate prediction by theory of normal force acting on slender body of rev- olution with curved profiles at angle-of-attack re- sults from inaccurate prediction of potential flow pressure distribution due to angle of attack, and from neglecting effect of cross-flow separation. In- crease in total axial force with angle of attack was primarily due to increase in base pressure force. NACA RM E51H23 WIRE CLOTH AS POROUS MATERIAL FOR TRANSPIRATION-COOLED WALLS. E. R. G. Eckert, Martin R. Kinsler, and Reeves P. Cochran. November 1951. 38p. diagrs., photos., tab. (NACA RM E51H23) (Declassified from Confidential, 9/15/55) The permeability characteristics and tensile strength of a porous material developed from stainless-steel corduroy wire cloth for use in transpiration-cooled vvalls where the primary stresses are in one direc- tion were investigated. The results of this investi- gation are presented and compared with similar re- sults obtained with porous sintered metal compacts. A much wider range of permeabilities is obtainable with the wire cloth than with the porous metal com- pacts considered and the ultimate tensile strength in the direction of the primary stresses for porous ma- terials produced from three mesh sizes of wire cloth is from two to three times the ultimate tensile strengths of the porous metal compacts. NACA RM E52I12 INTERSTAGE SURVEYS AND ANALYSIS OF VIS- COUS ACTION IN LATTER STAGES OF A MULTI- STAGE AXIAL-FLOW COMPRESSOR. William B. Briggs and Charles C. Giamati. March 1953. 51p. diagrs., photo., tab. (NACA RM E52I12) (Declassified from Confidential, 9/15/55) The overall performance of an eight-stage axial- flow compressor having a design stage pressure ratio of 1.23 was determined and radial interstage surveys were made for four weight flows at both 50 percent and 100 percent design speed. The survey data, from behind the guide vanes and each row of NACA RESEARCH ABSTRACTS NO. 91 the fifth to eighth stages, were presented as axial and tangential velocity components from which boundary-layer parameters of displacement and momentum thickness were calculated. An analysis related current qualitative ideas of flows which are imposed upon the through flow and the quantitative status of theory to the observed variation of these parameters. Midchannel velocity peaks in excess of design value were not found. NACA RM E51I25 FLOW SEPARATION AHEAD OF A BLUNT AXDALLY SYMMETRIC BODY AT MACH NUMBERS 1.76 TO 2.10. W. E. Moeckel. December 1951. 12p. diagrs., photos. (NACA RM E51I25) (Declassified from Confidential, 9/15/55) The pressure distribution and drag were determined for a spherical-nosed axially symmetric body with thin projecting rods at Mach numbers of 1.76, 1.93, and 2.10. The upstream projection distance of the rods was varied over a wide range to study changes in the character of the flow separation and to deter- mine the variation of drag and pressure distribution with tip projection. Drag coefficients between 0.18 and 0.30 were obtained for most tip projections at each Mach number. NACA RM L8I08 LONGITUDINAL-STABILITY INVESTIGATION OF HIGH-LIFT AND STALL-CONTROL DEVICES ON A 52° SWEPTBACK WING WITH AND WITHOUT FUSE- LAGE AND HORIZONTAL TAIL AT A REYNOLDS NUMBER OF 6.8 x 10 6 . Gerald V. Foster and James E. Fitzpatrick. December 20, 1948. 41p. diagrs., photos., tabs. (NACA RM L8I08) (Declassified from Confidential, 9/15/55) Contains low-speed longitudinal stability character- istics of a 52° sweptback wing of aspect ratio 2.88, taper ratio 0.625, and NACA 64j-112 airfoil sections normal to the 0.282-chord line, in combination with split flaps, leading-edge flaps, and upper-surface fences. Low-wing and midwing-fuselage aerody- namic characteristics are presented with and without a horizontal tail at various vertical locations. Tests were conducted at a Reynolds number of 6.8 x 10". NACA RM L8I30a EFFECTS OF A SWEPTBACK HYDROFOIL ON THE FORCE AND LONGITUDINAL STABILITY CHARAC- TERISTICS OF A TYPICAL HIGH-SPEED AIR- PLANE. Raymond B. Wood. December 2, 1948. 19p. diagrs., photo., tabs. (NACA RM L8I30a) (Declassified from Confidential, 9/15/55) An investigation was conducted in the Langley 8-foot high-speed tunnel to determine the effects of a swept- back hydrofoil on the force and longitudinal stability NACA RESEARCH ABSTRACTS NO. 9) characteristics of a typical high-speed airplane. The Mach number range for this investigation was from 0.60 to 0.95 and at M = 1.20. The effects of the hydrofoil on the lift, drag, and pitching-moment characteristics are presented. NACA RM L9I30 FLIGHT INVESTIGATIONS AT HIGH-SUBSONIC, TRANSONIC, AND SUPERSONIC SPEEDS TO DE- TERMINE ZERO-LIFT DRAG OF FIN-STABILIZED BODIES OF REVOLUTION HAVING FINENESS RATIOS OF 12.5, 8.91, AND 6.04 AND VARYING POSITIONS OF MAXIMUM DIAMETER. Roger G. Hart and Ellis R. Katz. November 30, 1949. 36p. diagrs., photos. (NACA RM L9I30) (Declassified from Confidential, 9/15/55) Rocket-powered models were flown at transonic and supersonic speeds to determine the zero-lift drag of fin-stabilized bodies of revolution differing in fine- ness ratio and in position of maximum diameter. The bodies were of fineness ratio 12.5, 8.91, and 6.04 and all had cut-off sterns with equal base and frontal areas. NACA RM L50I27 THE LONGITUDINAL STABILITY, CONTROL EF- FECTIVENESS, AND CONTROL HINGE-MOMENT CHARACTERISTICS OBTAINED FROM A FLIGHT INVESTIGATION OF A CANARD MISSILE CONFIG- URATION AT TRANSONIC AND SUPERSONIC Sl'EEDS. Roy J. Niewald and Martin T. Moul. November 24, 1950. 43p. diagrs., photos. (NACA RM L50I27) (Declassified from Confidential, 9/15/55) A 60° delta-wing canard missile configuration was flight-tested at the Langley Pilotless Aircraft Research Station at Wallops Island, Va. Longitudi- nal stability derivatives, control hinge -moment, and drag characteristics were obtained at transonic and supersonic velocities by utilizing a continuous step control disturbance of ±5°. NACA RM L51G31 SYSTEMATIC TWO-DIMENSIONAL CASCADE TESTS OF NACA 65-SERIES COMPRESSOR BLADES AT LOW SPEEDS. L. Joseph Herrig, James C. Emery, and John R. Erwin. September 1951. 223p. diagrs., photo., tabs. (NACA RM L51G31) (Declassified from Confidential, 9/15/55) A two-dimensional low-speed porous-wall cascade tunnel investigation has been conducted to establish the performance of the NACA 65-series compressor blade sections over the useful range of inlet angle, solidity, and section camber. Design points for opti- mum high-speed operation are presented. The load- ing limitation is determined for some conditions. Trends of section operating range with increasing section camber are determined for the four inlet angles tested. NACA RM L51H03 ROLLING EFFECTIVENESS OF ALL-MOVABLE WINGS AT SMALL ANGLES OF INCIDENCE AT MACH NUMBERS FROM 0.6 TO 1.6. H. Kurt Strass and Edward T. Marley. October 1951. 16p. diagrs., photo., tab. (NACA RM L51H03) (Declassified from Confidential, 9/15/55) An investigation of the rolling effectiveness of sev- eral all-movable wing configurations has been con- ducted throughout a Mach number range from 0.6 to 1.6 in order to check a simplified wing incidence cor- rection theory which states that for all-movable wing pb 2i w (\ + 2x\ , .... 5^ = * - h o ( i oy )• In addition, a comparison was made with two other more complex methods of esti- mation. The results showed that the simplified theory gave accurate agreement with experiment at Mach numbers from 0.6 to 1.6 and suggest the use of this simplified equation as a means of predicting the rolling effectiveness of all-movable wings throughout this speed range. NACA RM L52I16 SOME EXPERIMENTAL STUDIES OF PANEL FLUTTER AT MACH NUMBER 1.3. Maurice A. Sylvester and John E. Baker. December 1952. 25p. diagrs., photos., tab. (NACA RM L52I16) (Declassified from Confidential, 9/15/55) Experimental studies of panel flutter using thin metal plates .vere conducted at a Mach number of 1.3 to verify its existence and to study the effects of some structural parameters on the flutter characteristics. The effects of tensile forces and buckling were studied on panels clamped front and rear, in addition to initially buckled panels clamped on all four edges. Panel flutter was obtained under these laboratory conditions and it was found that tensile forces, shortening the panels, and increasing the bending stiffness were effective means for eliminating flutter. Buckled panels were more susceptible to flutter than unbuckled panels. No apparent systematic trends in the flutter modes or frequencies could be observed. NACA RM L53F17 METHOD OF ESTIMATING THE INCOMPRESSIBLE- FLOW PRESSURE DISTRIBUTION OF COMPRESSOR BLADE SECTIONS AT DESIGN ANGLE OF ATTACK. John R. Erwin and Laura A. Yacobi. December 1953. 41p. diagrs., tab. (NACA RM L53F17) (Declassified from Confidential, 9/15/55) A method was devised for estimating the incom- pressible flow pressure distribution over com- pressor blade sections at design angle of attack. The theoretical incremental velocities due to camber and thickness of the section as an isolated airfoil are assumed proportional to the average passage velocity and are modified by empirically determined interference factors. Comparisons were made be- tween estimated and test pressure distributions of NACA 65-series sections for typical conditions. Good agreement was obtained. 10 NACA RM L53I11 UNIVERSITY OF FLORIDA 3 1262 08153 283 9 NACA RESEARCH ABSTRACTS NO. 91 FREE-FLIGHT -TUNNEL INVESTIGATION OF THE LOW-SPEED STABILITY AND CONTROL CHARAC- TERISTICS OF A CANARD AIRPLANE MODEL. Joseph L. Johnson, Jr., and John W. Paulson. October 1953. 37p. diagrs., photo., 2 tabs. (NACA RM L53I11) (Declassified from Confidential, 9/15/55) Results are presented of an experimental investiga- tion in the Langley free-flight tunnel to determine the dynamic lateral stability and control character- istics of a model of a canard-type airplane. Tests were made with several vertical-tail configurations for the model with a triangular horizontal tail and with a sweptback horizontal tail having a leading- edge flap. NACA RM L53I25a FLIGHT DETERMINATION OF DRAG OF NORMAL- SHOCK NOSE INLETS WITH VARIOUS COWLING PROFILES AT MACH NUMBERS FROM 0.9 TO 1.5. R. I. Sears, C. F. Merlet, and L. W. Putland. October 1953. 36p. diagrs., photos., tabs. (NACA RM L53I25a) (Declassified from Confidential, 9/15/55) External-drag data are presented for normal-shock nose inlets with 1 -series, parabolic, and conic cowling profiles. The tests were made at an angle of attack of 0° by using rocket -propelled models in free flight at Mach numbers from 0.9 to 1.5. The Reynolds number based on body maximum diameter varied from 2.5 x 10° to 5.5 x 10°\ At maximum flow rate, the inlet models had about the same ex- ternal drag at a Mach number of approximately 1.1, but at higher Mach numbers the sharp-lip conic cowl had the least drag. Blunting or beveling the lip of the conic cowl while keeping the fineness ratio constant resulted in drag coefficients slightly higher than for the sharp-lip conic cowl at maximum flow rate. At a mass-flow ratio of about 0.8, the conic cowls with sharp, blunt, or beveled lips and the parabolic cowl all gave about the same drag. NACA RM L53I29b INVESTIGATION AT SUPERSONIC SPEEDS OF THE VARIATION WITH REYNOLDS NUMBER AND MACH NUMBER OF THE TOTAL, BASE, AND SKIN- FRICTION DRAG OF SEVEN BOATTAIL BODIES OF REVOLUTION DESIGNED FOR MINIMUM WAVE DRAG. August F. Bromm, Jr.. and Julia M. Goodwin. December 1953. 20p. diagrs., photo. (NACA RM L53I29b) (Declassified from Confidential, 9/15/55) Results are presented from an investigation of the variation with Reynolds number and Mach number of the total, base, and skin-friction drag of seven boat- tail bodies of revolution designed for minimum wave drag according to the theory of NACA TN 2550. The tests covered a Reynolds number range from approxi- mately 1.0 x 10 6 to 10.0 x 10 6 at Mach numbers of 1.62, 1.93, and 2.41, respectively. NACA RM L53I30b TWO-DIMENSIONAL LOW -SPEED CASCADE INVES- TIGATION OF NACA COMPRESSOR BLADE SEC- TIONS HAVING A SYSTEMATIC VARIATION IN MEAN-LINE LOADING. John R. Erwin, Melvyn Savage, and James C. Emery. November 1953. 129p. diagrs., tabs. (NACA RM L53I30b) (Declassified from Confidential, 9/15/55) The low-speed cascade performance of the high- speed NACA 65- (C-i A2l8b) 1( ^ compressor blade sections has been systematically investigated. When used in conjunction with published cascade data, the results will provide design information for all inlet angle and solidity conditions within the usual range of application. Summary curves have been prepared to facilitate the selection of blade sections and settings to fulfill the conditions dictated by compressor design velocity diagrams. Comparative tests of blade sections having widely different loading distributions indicated that these data, in conjunction with previ- ously published cascade data, permit a fairly accurate prediction of design performance for most compressor blade sections since the mean lines tested probably encompass the practical range of compressor-blade mean-line loading distributions. A comparative evaluation of the high-speed per- formance capabilities of the blade sections investi- gated was made. NACA - Langley Field, Va.